This application relates to an improved cooling circuit for a blade outer air seal, in which a plurality of distinct cooling schemes are utilized.
Gas turbine engines are provided with a number of functional sections, including a fan section, a compressor section, a combustion section, and a turbine section. Air and fuel are combusted in the combustion section. The products of the combustion move downstream, and pass over a series of turbine rotors, driving the rotors to create power.
It is desirable to have the bulk of the products of combustion pass over the turbine blade. Thus, a seal is placed circumferentially about the turbine rotors slightly radially spaced from a radially outer surface of the turbine blades. The seal is in a harsh environment, and must be able to withstand high temperatures. To address the high temperatures, the seal is typically provided with internal cooling channels. Air circulates through the cooling channels to cool the seal.
In the prior art, one type of cooling scheme has been utilized across the seal. However, the cooling challenges faced across the seal vary. As an example, the seal extends from a leading edge to a trailing edge. A pressure ratio between the cooling air and the working air is low at the leading edge, and greater at the trailing edge. Even so, the prior art has not tailored the cooling channels to the location. Further, the prior art has typically used only relatively large cooling channels in the blade outer air seals.
More recently, compact heat exchanger cooling schemes (or microcircuit cooling channels) have been developed, which utilize relatively thin and small passages to convey cooling air through a body. These compact heat exchangers are formed by lost core molding techniques. While these techniques provide efficient and effective cooling, they have not been applied to a cool blade outer air seal.